XFOIL         Version 6.94
  
 Calculated polar for: FX 66-S-196 AIRFOIL                             
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.080     Re =     0.450 e 6     Ncrit =   9.000
  
  alpha     CL        CD       CDp       CM    Top_Xtr Bot_Xtr
 ------- -------- --------- --------- -------- ------- -------
   0.500   0.5238   0.01273   0.00615  -0.1005  0.5821  0.6167
   1.000   0.5818   0.01291   0.00627  -0.1016  0.5792  0.6176
   1.500   0.6385   0.01323   0.00657  -0.1025  0.5759  0.6185
   2.000   0.6941   0.01329   0.00668  -0.1032  0.5736  0.6197
   2.500   0.7493   0.01340   0.00686  -0.1038  0.5710  0.6212
   3.000   0.8044   0.01358   0.00712  -0.1043  0.5682  0.6233
   3.500   0.8599   0.01376   0.00736  -0.1049  0.5651  0.6255
   4.000   0.9165   0.01388   0.00752  -0.1057  0.5616  0.6282
   4.500   0.9735   0.01415   0.00783  -0.1066  0.5587  0.6310
   5.000   1.0279   0.01478   0.00854  -0.1072  0.5548  0.6338
   5.500   1.0788   0.01493   0.00885  -0.1070  0.5518  0.6365
   6.000   1.1292   0.01519   0.00925  -0.1067  0.5479  0.6393
   6.500   1.1808   0.01541   0.00963  -0.1067  0.5441  0.6433
   7.000   1.2371   0.01532   0.00965  -0.1073  0.5392  0.6482
   7.500   1.2898   0.01528   0.00970  -0.1072  0.5315  0.6541
   8.000   1.3348   0.01468   0.00920  -0.1053  0.5200  0.6609
   8.500   1.3765   0.01441   0.00908  -0.1030  0.5082  0.6698
   9.000   1.4082   0.01418   0.00891  -0.0988  0.4924  0.6809
   9.500   1.4129   0.01434   0.00934  -0.0897  0.4750  0.6949
  10.000   1.4199   0.01497   0.01012  -0.0821  0.4542  0.7141
  10.500   1.4119   0.01646   0.01172  -0.0731  0.4228  0.7428
  11.000   1.3749   0.01971   0.01504  -0.0618  0.3762  0.8055